This invention relates generally to gas turbine engines and, more particularly, to gas turbine engine compressors.
At least some known gas turbine engines include a multi-stage axial compressor, a combustor, and a turbine. Airflow entering the compressor is compressed and channeled towards the combustor wherein the airflow is mixed with fuel and ignited, producing hot combustion gases used to drive the turbine. At least one known gas turbine engine includes a High Pressure Centrifugal Compressor (HPCC) that operates by inducing a centrifugal force to an air mass to achieve compression. Specifically, in at least some known gas turbine engines, the Centrifugal Compressor includes an impeller that is configured to add energy to the compressor and a diffusing system that is configured to convert a kinetic portion of the added energy into static pressure. In at least some known Centrifugal Compressors, the diffuser includes a radial diffuser, a bend, and a deswirler. In some known Centrifugal Compressors the radial diffuser, the bend, and the deswirler are made as an integral part.
At least one known gas turbine engine determines a centrifugal stage pressure ratio based on the impeller tip speed and basic geometric parameters, i.e., the blade exit, impeller tip height, back-sweep, the impeller inlet and exit radii, and an estimate of the impeller hub axial length. The maximum pressure ratio of known centrifugal compressors is generally limited by the highest tip speed allowed by its material properties and stall margins. For higher pressure ratios, known compressors use rearward-swept blades at the impeller exit to facilitate enhanced stall margin and operating efficiency. Specifically, to increasing compressor pressure ratio may require increasing both impeller tip speed and back-sweep to facilitate alleviating an impeller blade aerodynamic loading “diffusion”, such that an efficiency is enhanced and a sufficient stall margin is secured.